In gas turbine engines for use in powering aircraft, air is directed through multiple stage compressors as it flows axially or axially and radially through the engine to a combuster. As the air passes through each successive compressor stage, the pressure of the air is increased. Under certain conditions, such as when the engine is operating at off design conditions, interstage bleed is required to match the compressor stages. If this compressor matching is not acheived an engine surge or blow-out may occur, endangering the operation of the engine and the associated aircraft.
To mitigate against these conditions, such gas turbine engines have incorporated bleed valves in the engine casing forward of the burner which, when an engine surge is imminent, open to rematch the compressor stages. These bleed valves have taken many forms from simple ports in the compressor casing which open via a movable valve element to devices which separate adjacent segments of the engine casing thereby creating an opening there between.
However, these valves, although useful, present problems where the air bleed off is directed into a secondary air flow, in lieu of being dumped overboard. In the design of these prior art bleed valves all of the criteria which must be met such as, simple maintenance of the valve, maintenance of a smooth fluid flow through the bypass flow path and quick response time are not all addressed in any single prior art valve.
Therefore, what is necessary in this art is a bleed valve that is simple to service, minimizes the disturbance to the secondary air flow and offers quick response to the pressure changes which lead to the engine operating problems.